This invention relates to internal cooling within a gas turbine engine; and more particularly, to apparatus and method for providing better and more uniform cooling in a transition region between a combustion section and discharge section of the turbine.
Traditional gas turbine combustors use diffusion (i.e., non-premixed) combustion in which fuel and air enter the combustion chamber separately. The process of mixing and burning produces flame temperatures exceeding 3900° F. Since conventional combustors and/or transition pieces having liners are generally capable of withstanding a maximum temperature on the order of only about 1500° F. for about ten thousand hours (10,000 hrs.), steps to protect the combustor and/or transition piece must be taken. This has typically been done by film-cooling which involves introducing relatively cool compressor air into a plenum formed by the combustor liner surrounding the outside of the combustor. In this prior arrangement, the air from the plenum passes through louvers in the combustor liner and then passes as a film over the inner surface of the liner, thereby maintaining combustor liner integrity.
Because diatomic nitrogen rapidly disassociates at temperatures exceeding about 3000° F. (about 1650° C.), the high temperatures of diffusion combustion result in relatively large NOx emissions. One approach to reducing NOx emissions has been to premix the maximum possible amount of compressor air with fuel. The resulting lean premixed combustion produces cooler flame temperatures and thus lower NOx emissions. Although lean premixed combustion is cooler than diffusion combustion, the flame temperature is still too hot for prior conventional combustor components to withstand.
Furthermore, because the advanced combustors premix the maximum possible amount of air with the fuel for NOx reduction, little or no cooling air is available, making film-cooling of the combustor liner and transition piece premature at best. Nevertheless, combustor liners require active cooling to maintain material temperatures below limits. In dry low NOx (DLN) emission systems, this cooling can only be supplied as cold side convection. Such cooling must be performed within the requirements of thermal gradients and pressure loss. Thus, means such as thermal barrier coatings in conjunction with “backside” cooling have been considered to protect the combustor liner and transition piece from destruction by such high heat. Backside cooling involved passing the compressor discharge air over the outer surface of the transition piece and combustor liner prior to premixing the air with the fuel.
With respect to the combustor liner, one current practice is to impingement cool the liner, or to provide linear turbulators on the exterior surface of the liner. Another more recent practice is to provide an array of concavities on the exterior or outside surface of the liner (see U.S. Pat. No. 6,098,397). The various known techniques enhance heat transfer but with varying effects on thermal gradients and pressure losses. Turbulation strips work by providing a blunt body in the flow which disrupts the flow creating shear layers and high turbulence to enhance heat transfer on the surface. Dimple concavities function by providing organized vortices that enhance flow mixing and scrub the surface to improve heat transfer.
A low heat transfer rate from the liner can lead to high liner surface temperatures and ultimately loss of strength. Several potential failure modes due to the high temperature of the liner include, but are not limited to, cracking of the aft sleeve weld line, bulging and triangulation. These mechanisms shorten the life of the liner, requiring replacement of the part prematurely.
Accordingly, there remains a need for enhanced levels of active cooling with minimal pressure losses at higher firing temperatures than previously available while extending a combustion inspection interval to decrease the cost to produce electricity.